US20140050558A1 - Temperature gradient management arrangement for a turbine system and method of managing a temperature gradient of a turbine system - Google Patents
Temperature gradient management arrangement for a turbine system and method of managing a temperature gradient of a turbine system Download PDFInfo
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- US20140050558A1 US20140050558A1 US13/586,354 US201213586354A US2014050558A1 US 20140050558 A1 US20140050558 A1 US 20140050558A1 US 201213586354 A US201213586354 A US 201213586354A US 2014050558 A1 US2014050558 A1 US 2014050558A1
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- turbine system
- temperature gradient
- outlet
- management arrangement
- flow rate
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- 230000001965 increasing effect Effects 0.000 claims description 3
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- 239000000446 fuel Substances 0.000 description 8
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- 229910052739 hydrogen Inorganic materials 0.000 description 1
- 230000001939 inductive effect Effects 0.000 description 1
- 239000007788 liquid Substances 0.000 description 1
- 239000003345 natural gas Substances 0.000 description 1
- 238000010926 purge Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 230000008646 thermal stress Effects 0.000 description 1
- 230000007704 transition Effects 0.000 description 1
Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
- F02C6/04—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output
- F02C6/06—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
- F02C6/08—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/025—Seal clearance control; Floating assembly; Adaptation means to differential thermal dilatations
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
- F01D11/06—Control thereof
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
- F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
- F01D5/082—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades on the side of the rotor disc
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
- F04D27/02—Surge control
- F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
- F04D27/0223—Control schemes therefor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/05—Shafts or bearings, or assemblies thereof, specially adapted for elastic fluid pumps
- F04D29/053—Shafts
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/58—Cooling; Heating; Diminishing heat transfer
- F04D29/582—Cooling; Heating; Diminishing heat transfer specially adapted for elastic fluid pumps
- F04D29/584—Cooling; Heating; Diminishing heat transfer specially adapted for elastic fluid pumps cooling or heating the machine
Definitions
- the subject matter disclosed herein relates to turbine systems, and more particularly to a temperature gradient management arrangement for such turbine systems, as well as a method of managing temperature gradients.
- Turbine systems typically include a rotor having a plurality of stacked wheels.
- An outer radial region of the stacked wheels is known as a rim portion while a central radial region of the stacked wheels is known as a bore portion.
- Typical operation of turbine systems entails high temperatures which may subject various components of the turbine to relatively extreme thermal loads.
- the high temperatures of the main flow path are the result of the compression process and the combustion process within the turbine system.
- large temperature gradients in the stacked wheels are often present, particularly during startup of the turbine system while main flow path temperatures are significantly higher in temperature than the rotor.
- the large temperature gradients are caused by the rim portion being significantly hotter than the bore portion.
- the large temperature gradients described above impart thermal stresses that are superimposed on the mechanical stresses due to centrifugal forces and surface pressures.
- the stress and temperature history experienced by the rotor components determines damage accumulated over each operating cycle and therefore the life expectancy of the rotor.
- a temperature gradient management arrangement for a turbine system includes a rotor comprising a rotor bore extending axially along the rotor. Also included is a secondary flow path comprising an inlet for a secondary airflow to flow to the rotor bore and an outlet disposed axially upstream of the inlet, relative to a main flow direction of the turbine system. Further included is a flow rate manipulator disposed proximate the outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
- a temperature gradient management arrangement for a turbine system includes a rotor comprising a plurality of stacked wheels and a rotor bore. Also included is a plurality of compressor stages. Further included is at least one secondary flow path comprising at least one inlet for a secondary airflow to flow to the rotor bore, wherein the at least one secondary flow path extends from the at least one inlet in an upstream direction relative to a main flow direction toward at least one outlet disposed proximate at least one of the plurality of compressor stages. Yet further included is a flow rate manipulator disposed proximate the at least one outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
- a method of managing a temperature gradient of a turbine system includes routing a secondary airflow through an inlet to a rotor bore disposed within a rotor along a secondary flow path to an outlet. Also included is increasing a flow path clearance at the outlet to increase a flow rate of the secondary airflow through the secondary flow path during a first turbine system operating condition. Further included is decreasing the flow path clearance at the outlet to decrease the flow rate of the secondary airflow through the secondary flow path during a second turbine system operating condition.
- FIG. 1 is a schematic illustration of a turbine system
- FIG. 2 is a partial, schematic side view of the turbine system
- FIG. 3 is a flow diagram illustrating a method of managing a temperature gradient of the turbine system.
- axial and axially refer to directions and orientations extending substantially parallel to a center longitudinal axis of a turbine system.
- radial and radially refer to directions and orientations extending substantially orthogonally to the center longitudinal axis of the turbine system.
- upstream and downstream refer to directions and orientations relative to an axial flow direction with respect to the center longitudinal axis of the turbine system.
- a turbine system such as a gas turbine system, for example, is schematically illustrated with reference numeral 10 .
- the gas turbine system 10 includes a compressor section 12 , a combustor section 14 , a turbine section 16 , a rotor 18 and a fuel nozzle 20 .
- one embodiment of the gas turbine system 10 may include a plurality of compressors 12 , combustors 14 , turbines 16 , rotors 18 and fuel nozzles 20 .
- the compressor section 12 and the turbine section 16 are coupled by the rotor 18 .
- the rotor 18 comprises a single shaft or a plurality of shaft segments coupled together to form the rotor 18 .
- the combustor section 14 uses a combustible liquid and/or gas fuel, such as natural gas or a hydrogen rich synthetic gas, to run the gas turbine system 10 .
- fuel nozzles 20 are in fluid communication with a main flow path 26 exiting the compressor 12 and a fuel supply 22 .
- the fuel nozzles 20 create an air-fuel mixture, and discharge the air-fuel mixture into the combustor section 14 , thereby causing a combustion that creates a hot pressurized exhaust gas 28 .
- the combustor section 14 directs the hot pressurized gas through a transition piece into a turbine nozzle (or “stage one nozzle”), and other stages of buckets and nozzles causing rotation of turbine blades within an outer casing 24 of the turbine section 16 .
- a plurality of stacked wheels includes a plurality of solid wheels 30 and a plurality of annular wheels 32 , with the plurality of solid wheels 30 arranged alternately between the plurality of annular wheels 32 .
- Both the plurality of solid wheels 30 and the plurality of annular wheels 32 are mounted on, and form, a portion of the rotor 18 .
- the rotor 18 includes a rim portion 34 disposed at a radially outward position of the rotor 18 , as well as a rotor bore 36 extending relatively axially throughout the rotor 18 , however, it is to be appreciated that the rotor bore 36 may include a tortuous path that extends in radial and/or circumferential directions.
- Each of the plurality of solid wheels 30 and the plurality of annular wheels 32 includes a rotor blade 38 projecting radially outwardly from the rotor 18 , while a plurality of stator vanes 40 are mounted on a stator (not illustrated).
- Each of the plurality of stator vanes 40 is typically positioned alternately between the rotor blades 38 and for illustration simplicity, only two of the plurality of stator vanes 40 are referenced.
- the rotor blades 38 and the plurality of stator vanes 40 form a passage through which the main flow path 26 in the compressor section 12 flows. The temperature of the main flow path 26 increases progressively as its pressure increases through the compressor section 12 .
- the rim portion 34 of the rotor 18 is exposed to the progressively hotter main flow path 26 , while the rotor bore 36 remains shielded from the main flow path 26 .
- the rotor bore 36 forms the root of the rotor 18 , thereby extending proximate a rotor centerline 42 .
- a temperature gradient often exists between the rim portion 34 and the rotor bore 36 of the rotor 18 .
- the rotor bore 36 may be heated during startup of the gas turbine system 10 and cooled during shutdown, which are the two operating periods that typically result in particularly high temperature gradients.
- the compressor section 12 includes a plurality of compressor stages 44 , with each stage comprising one or more circumferentially spaced stator vanes aligned in a row at a similar axial location, along with an axially preceding or succeeding row of circumferentially spaced rotor blades disposed at a similar axial location.
- the plurality of compressor stages 44 include a middle stage 46 disposed at a relatively axial mid-point of the plurality of compressor stages 44 .
- a plurality of forward stages 48 are positioned upstream of the middle stage 46 , with respect to a direction of the main flow path 26 flowing through the compressor section 12 .
- a plurality of aft stages 50 are positioned downstream of the middle stage 46 , also with respect to a direction of the main flow path 26 .
- heating or cooling of the rotor bore 36 may beneficially reduce the temperature gradient present between radially inner portions of the rotor 18 , such as the rotor bore 36 itself, and radially outer portions of the rotor 18 , such as the rim portion 34 . Accordingly, a secondary airflow 52 is provided from the main flow path 26 flowing throughout the compressor section 12 to the rotor bore 36 . The remainder of the main flow path 26 typically flows to the combustor section 14 and as a cool/purge flow 53 for the turbine section 16 .
- the secondary airflow 52 is routed to the rotor bore 36 through at least one inlet 54 disposed proximate the middle stage 46 and/or at least one of the plurality of aft stages 50 , with the at least one inlet 54 part of a secondary flow path 56 . Subsequent to routing of the secondary airflow 52 through the at least one inlet 54 , the secondary airflow 52 is directed upstream (relative to the main flow path 26 ) along the secondary flow path 56 that is defined by the rotor bore 36 . As the secondary airflow 52 travels upstream, the rotor bore 36 , and therefore the radially inner portion of the rotor 18 are heated during startup to reduce the temperature gradient between the rotor bore 36 and the rim portion 34 .
- the secondary flow path 56 that the secondary airflow 52 is routed along may be of various path dimensions and shapes.
- the secondary flow path 56 may extend around the annular wheels 32 and through the solid wheels 30 , thereby forming a curved flow path referred to as a serpentine flow path.
- the secondary flow path 56 typically extends upstream from the at least one inlet 54 to at least one outlet 58 disposed proximate at least one of the plurality of forward stages 48 .
- the at least one outlet 58 may extend through a variety of components proximate at least one of the plurality of forward stages 48 , it is contemplated that the at least one outlet 58 is disposed at a stator vane diaphragm 60 proximate the rim portion 34 of the rotor 18 .
- a flow rate manipulator 62 is located within or proximate the at least one outlet 58 to control a clearance 64 that facilitates expulsion of the secondary airflow 52 from the secondary flow path 56 .
- the flow rate manipulator 62 comprises an adjustable seal that alters the flow rate of the secondary airflow 52 traveling throughout the secondary flow path 56 .
- the rotor components may undergo an axial deflection with respect to the stator parts due to a combination of temperature differential, associated thermal expansion and load stresses. The transiently varying relative deflection reaches a maximum during the full load operating time and remains constant during the entire duration of steady state, or near steady state, operation of the gas turbine system 10 .
- the flow rate manipulator 62 may be positioned such that it does not block the at least one outlet 58 , thereby inducing the secondary airflow 52 throughout the secondary flow path 56 due to a pressure differential between the rotor bore 36 and the main flow path 26 .
- the flow rate manipulator may relatively move and begin to cover and block at least a portion of the at least one outlet 58 , thereby restricting the flow of the secondary airflow 52 by decreasing the clearance 64 , which in turn decreases the flow rate of the secondary airflow 52 within the secondary flow path 56 for heating therealong.
- the flow rate manipulator 62 slows or stops movement, with respect to the rotor components and may continue to cover all or part of the at least one outlet 58 .
- the method of managing a temperature gradient of a turbine system 100 includes routing a secondary airflow through at least one inlet to a rotor bore 102 disposed within the rotor 18 along the secondary flow path 56 to the at least one outlet 58 .
- a flow path clearance is increased at the at least one outlet during a first turbine system operating condition 104 in order to increase a flow rate of the secondary airflow 52 .
- the flow path clearance is decreased at the at least one outlet during a second turbine operating condition 106 in order to decrease the flow rate of the secondary airflow 52 .
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Life Sciences & Earth Sciences (AREA)
- Sustainable Development (AREA)
- Physics & Mathematics (AREA)
- Thermal Sciences (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A temperature gradient management arrangement for a turbine system includes a rotor comprising a rotor bore extending axially along the rotor. Also included is a secondary flow path comprising an inlet for a secondary airflow to flow to the rotor bore and an outlet disposed axially upstream of the inlet, relative to a main flow direction of the turbine system. Further included is a flow rate manipulator disposed proximate the outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
Description
- The subject matter disclosed herein relates to turbine systems, and more particularly to a temperature gradient management arrangement for such turbine systems, as well as a method of managing temperature gradients.
- Turbine systems typically include a rotor having a plurality of stacked wheels. An outer radial region of the stacked wheels is known as a rim portion while a central radial region of the stacked wheels is known as a bore portion. Typical operation of turbine systems entails high temperatures which may subject various components of the turbine to relatively extreme thermal loads. The high temperatures of the main flow path are the result of the compression process and the combustion process within the turbine system. As a result, large temperature gradients in the stacked wheels are often present, particularly during startup of the turbine system while main flow path temperatures are significantly higher in temperature than the rotor. Specifically, the large temperature gradients are caused by the rim portion being significantly hotter than the bore portion.
- The large temperature gradients described above impart thermal stresses that are superimposed on the mechanical stresses due to centrifugal forces and surface pressures. The stress and temperature history experienced by the rotor components determines damage accumulated over each operating cycle and therefore the life expectancy of the rotor.
- According to one aspect of the invention, a temperature gradient management arrangement for a turbine system includes a rotor comprising a rotor bore extending axially along the rotor. Also included is a secondary flow path comprising an inlet for a secondary airflow to flow to the rotor bore and an outlet disposed axially upstream of the inlet, relative to a main flow direction of the turbine system. Further included is a flow rate manipulator disposed proximate the outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
- According to another aspect of the invention, a temperature gradient management arrangement for a turbine system includes a rotor comprising a plurality of stacked wheels and a rotor bore. Also included is a plurality of compressor stages. Further included is at least one secondary flow path comprising at least one inlet for a secondary airflow to flow to the rotor bore, wherein the at least one secondary flow path extends from the at least one inlet in an upstream direction relative to a main flow direction toward at least one outlet disposed proximate at least one of the plurality of compressor stages. Yet further included is a flow rate manipulator disposed proximate the at least one outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
- According to yet another aspect of the invention, a method of managing a temperature gradient of a turbine system is provided. The method includes routing a secondary airflow through an inlet to a rotor bore disposed within a rotor along a secondary flow path to an outlet. Also included is increasing a flow path clearance at the outlet to increase a flow rate of the secondary airflow through the secondary flow path during a first turbine system operating condition. Further included is decreasing the flow path clearance at the outlet to decrease the flow rate of the secondary airflow through the secondary flow path during a second turbine system operating condition.
- These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
- The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:
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FIG. 1 is a schematic illustration of a turbine system; -
FIG. 2 is a partial, schematic side view of the turbine system; and -
FIG. 3 is a flow diagram illustrating a method of managing a temperature gradient of the turbine system. - The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings.
- The terms “axial” and “axially” as used in this application refer to directions and orientations extending substantially parallel to a center longitudinal axis of a turbine system. The terms “radial” and “radially” as used in this application refer to directions and orientations extending substantially orthogonally to the center longitudinal axis of the turbine system. The terms “upstream” and “downstream” as used in this application refer to directions and orientations relative to an axial flow direction with respect to the center longitudinal axis of the turbine system.
- Referring to
FIG. 1 , a turbine system, such as a gas turbine system, for example, is schematically illustrated withreference numeral 10. Thegas turbine system 10 includes acompressor section 12, acombustor section 14, aturbine section 16, arotor 18 and afuel nozzle 20. It is to be appreciated that one embodiment of thegas turbine system 10 may include a plurality ofcompressors 12,combustors 14,turbines 16,rotors 18 andfuel nozzles 20. Thecompressor section 12 and theturbine section 16 are coupled by therotor 18. Therotor 18 comprises a single shaft or a plurality of shaft segments coupled together to form therotor 18. - The
combustor section 14 uses a combustible liquid and/or gas fuel, such as natural gas or a hydrogen rich synthetic gas, to run thegas turbine system 10. For example,fuel nozzles 20 are in fluid communication with amain flow path 26 exiting thecompressor 12 and a fuel supply 22. Thefuel nozzles 20 create an air-fuel mixture, and discharge the air-fuel mixture into thecombustor section 14, thereby causing a combustion that creates a hot pressurizedexhaust gas 28. Thecombustor section 14 directs the hot pressurized gas through a transition piece into a turbine nozzle (or “stage one nozzle”), and other stages of buckets and nozzles causing rotation of turbine blades within anouter casing 24 of theturbine section 16. - Referring now to
FIG. 2 , a partial schematic illustrates in greater detail thecompressor section 12 and theturbine section 16, which are operably coupled by therotor 18. A plurality of stacked wheels includes a plurality ofsolid wheels 30 and a plurality ofannular wheels 32, with the plurality ofsolid wheels 30 arranged alternately between the plurality ofannular wheels 32. Both the plurality ofsolid wheels 30 and the plurality ofannular wheels 32 are mounted on, and form, a portion of therotor 18. Therotor 18 includes arim portion 34 disposed at a radially outward position of therotor 18, as well as a rotor bore 36 extending relatively axially throughout therotor 18, however, it is to be appreciated that the rotor bore 36 may include a tortuous path that extends in radial and/or circumferential directions. - Each of the plurality of
solid wheels 30 and the plurality ofannular wheels 32 includes arotor blade 38 projecting radially outwardly from therotor 18, while a plurality ofstator vanes 40 are mounted on a stator (not illustrated). Each of the plurality ofstator vanes 40 is typically positioned alternately between therotor blades 38 and for illustration simplicity, only two of the plurality ofstator vanes 40 are referenced. Therotor blades 38 and the plurality of stator vanes 40 form a passage through which themain flow path 26 in thecompressor section 12 flows. The temperature of themain flow path 26 increases progressively as its pressure increases through thecompressor section 12. Consequently, therim portion 34 of therotor 18 is exposed to the progressively hottermain flow path 26, while the rotor bore 36 remains shielded from themain flow path 26. The rotor bore 36 forms the root of therotor 18, thereby extending proximate arotor centerline 42. Based on exposure of therim portion 34 to themain flow path 26, a temperature gradient often exists between therim portion 34 and the rotor bore 36 of therotor 18. To mitigate the temperature gradient noted above, the rotor bore 36 may be heated during startup of thegas turbine system 10 and cooled during shutdown, which are the two operating periods that typically result in particularly high temperature gradients. - As illustrated, the
compressor section 12 includes a plurality ofcompressor stages 44, with each stage comprising one or more circumferentially spaced stator vanes aligned in a row at a similar axial location, along with an axially preceding or succeeding row of circumferentially spaced rotor blades disposed at a similar axial location. The plurality ofcompressor stages 44 include amiddle stage 46 disposed at a relatively axial mid-point of the plurality ofcompressor stages 44. A plurality offorward stages 48 are positioned upstream of themiddle stage 46, with respect to a direction of themain flow path 26 flowing through thecompressor section 12. Additionally, a plurality ofaft stages 50 are positioned downstream of themiddle stage 46, also with respect to a direction of themain flow path 26. - As described above, heating or cooling of the rotor bore 36 may beneficially reduce the temperature gradient present between radially inner portions of the
rotor 18, such as the rotor bore 36 itself, and radially outer portions of therotor 18, such as therim portion 34. Accordingly, asecondary airflow 52 is provided from themain flow path 26 flowing throughout thecompressor section 12 to the rotor bore 36. The remainder of themain flow path 26 typically flows to thecombustor section 14 and as a cool/purge flow 53 for theturbine section 16. Thesecondary airflow 52 is routed to the rotor bore 36 through at least oneinlet 54 disposed proximate themiddle stage 46 and/or at least one of the plurality ofaft stages 50, with the at least oneinlet 54 part of a secondary flow path 56. Subsequent to routing of thesecondary airflow 52 through the at least oneinlet 54, thesecondary airflow 52 is directed upstream (relative to the main flow path 26) along the secondary flow path 56 that is defined by the rotor bore 36. As thesecondary airflow 52 travels upstream, the rotor bore 36, and therefore the radially inner portion of therotor 18 are heated during startup to reduce the temperature gradient between the rotor bore 36 and therim portion 34. It is to be appreciated that the secondary flow path 56 that thesecondary airflow 52 is routed along may be of various path dimensions and shapes. For example, the secondary flow path 56 may extend around theannular wheels 32 and through thesolid wheels 30, thereby forming a curved flow path referred to as a serpentine flow path. - Irrespective of the precise dimensions and shape of the secondary flow path 56, it is to be understood that the secondary flow path 56 typically extends upstream from the at least one
inlet 54 to at least oneoutlet 58 disposed proximate at least one of the plurality of forward stages 48. Although the at least oneoutlet 58 may extend through a variety of components proximate at least one of the plurality of forward stages 48, it is contemplated that the at least oneoutlet 58 is disposed at astator vane diaphragm 60 proximate therim portion 34 of therotor 18. - A
flow rate manipulator 62 is located within or proximate the at least oneoutlet 58 to control aclearance 64 that facilitates expulsion of thesecondary airflow 52 from the secondary flow path 56. In one embodiment, theflow rate manipulator 62 comprises an adjustable seal that alters the flow rate of thesecondary airflow 52 traveling throughout the secondary flow path 56. Typically, during startup and shutdown of thegas turbine system 10, the rotor components may undergo an axial deflection with respect to the stator parts due to a combination of temperature differential, associated thermal expansion and load stresses. The transiently varying relative deflection reaches a maximum during the full load operating time and remains constant during the entire duration of steady state, or near steady state, operation of thegas turbine system 10. Thus, prior to and during startup, theflow rate manipulator 62 may be positioned such that it does not block the at least oneoutlet 58, thereby inducing thesecondary airflow 52 throughout the secondary flow path 56 due to a pressure differential between the rotor bore 36 and themain flow path 26. As the stator vane components to which theflow rate manipulator 62 is operably coupled to expand axially during transient state operation (referred to as a first operating condition) with respect to the rotor components, the flow rate manipulator may relatively move and begin to cover and block at least a portion of the at least oneoutlet 58, thereby restricting the flow of thesecondary airflow 52 by decreasing theclearance 64, which in turn decreases the flow rate of thesecondary airflow 52 within the secondary flow path 56 for heating therealong. As thegas turbine system 10 reaches steady state, or near steady state, operation, referred to as a second operating condition, theflow rate manipulator 62 slows or stops movement, with respect to the rotor components and may continue to cover all or part of the at least oneoutlet 58. - It is contemplated that various alternative arrangements may be employed to suitably increase and decrease the flow rate of the
secondary airflow 52 during transient state operation and steady state operation, respectively. - As illustrated in the flow diagram of
FIG. 3 , and with reference toFIGS. 1 and 2 , a method of managing a temperature gradient of aturbine system 100 is also provided. Thegas turbine system 10, and more specifically thecompressor section 12 androtor 18 have been previously described and specific structural components need not be described in further detail. The method of managing a temperature gradient of aturbine system 100 includes routing a secondary airflow through at least one inlet to arotor bore 102 disposed within therotor 18 along the secondary flow path 56 to the at least oneoutlet 58. A flow path clearance is increased at the at least one outlet during a first turbinesystem operating condition 104 in order to increase a flow rate of thesecondary airflow 52. Conversely, the flow path clearance is decreased at the at least one outlet during a secondturbine operating condition 106 in order to decrease the flow rate of thesecondary airflow 52. - While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.
Claims (20)
1. A temperature gradient management arrangement for a turbine system comprising:
a rotor comprising a rotor bore extending axially along the rotor;
a secondary flow path comprising an inlet for a secondary airflow to flow to the rotor bore and an outlet disposed axially upstream of the inlet, relative to a main flow direction of the turbine system; and
a flow rate manipulator disposed proximate the outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
2. The temperature gradient management arrangement of claim 1 , wherein the first turbine system operating condition comprises a transient state operation.
3. The temperature gradient management arrangement of claim 1 , wherein the second turbine system operating condition comprises a steady state operation.
4. The temperature gradient management arrangement of claim 1 , wherein the flow rate manipulator comprises at least one adjustable seal configured to control a clearance proximate the outlet.
5. The temperature gradient management arrangement of claim 4 , wherein the flow rate manipulator is disposed proximate a stator vane diaphragm.
6. The temperature gradient management arrangement of claim 1 , further comprising a plurality of compressor stages, the plurality of compressor stages comprising a middle stage, a plurality of forward stages disposed upstream of the middle stage and a plurality of aft stages disposed downstream of the middle stage.
7. The temperature gradient management arrangement of claim 6 , wherein the inlet is disposed proximate at least one of the middle stage and the plurality of aft stages.
8. The temperature gradient management arrangement of claim 7 , further comprising a plurality of inlets.
9. The temperature gradient management arrangement of claim 6 , wherein the outlet is disposed proximate at least one of the plurality of forward stages.
10. The temperature gradient management arrangement of claim 9 , further comprising a plurality of outlets.
11. A temperature gradient management arrangement for a turbine system comprising:
a rotor comprising a plurality of stacked wheels and a rotor bore;
a plurality of compressor stages;
at least one secondary flow path comprising at least one inlet for a secondary airflow to flow to the rotor bore, wherein the at least one secondary flow path extends from the at least one inlet in an upstream direction relative to the main flow toward at least one outlet disposed proximate at least one of the plurality of compressor stages; and
a flow rate manipulator disposed proximate the at least one outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
12. The temperature gradient management arrangement of claim 11 , wherein the first turbine system operating condition comprises a transient state operation.
13. The temperature gradient management arrangement of claim 11 , wherein the second turbine system operating condition comprises a steady state operation.
14. The temperature gradient management arrangement of claim 11 , wherein the flow rate manipulator comprises at least one adjustable seal configured to control a clearance proximate the at least one outlet.
15. The temperature gradient management arrangement of claim 14 , wherein the flow rate manipulator is disposed proximate a stator vane diaphragm.
16. The temperature gradient management arrangement of claim 11 , wherein the plurality of compressor stages comprises:
a middle stage;
a plurality of forward stages disposed upstream of the middle stage;
a plurality of aft stages disposed downstream of the middle stage; and
wherein the at least one inlet is disposed proximate at least one of the middle stage and the plurality of aft stages; and
wherein the at least one outlet is disposed proximate at least one of the plurality of forward stages.
17. A method of managing a temperature gradient of a turbine system comprising:
routing a secondary airflow through an inlet to a rotor bore disposed within a rotor along a secondary flow path to an outlet;
increasing a flow path clearance at the outlet to increase a flow rate of the secondary airflow through the secondary flow path during a first turbine system operating condition; and
decreasing the flow path clearance at the outlet to decrease the flow rate of the secondary airflow through the flow path during a second turbine system operating condition.
18. The method of claim 17 , wherein routing the secondary airflow further comprises directing the secondary airflow upstream along the secondary flow path to the outlet, wherein the outlet is disposed proximate a stator vane diaphragm.
19. The method of claim 17 , wherein the first turbine system operating condition comprises a transient state operation.
20. The method of claim 17 , wherein the second turbine system operating condition comprises a steady state operation.
Priority Applications (5)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/586,354 US20140050558A1 (en) | 2012-08-15 | 2012-08-15 | Temperature gradient management arrangement for a turbine system and method of managing a temperature gradient of a turbine system |
JP2013165704A JP2014037831A (en) | 2012-08-15 | 2013-08-09 | Temperature gradient management arrangement for turbine system and method of managing temperature gradient of turbine system |
DE102013108741.8A DE102013108741A1 (en) | 2012-08-15 | 2013-08-12 | A temperature gradient coping assembly for a turbine system and method for overcoming a temperature gradient of a turbine system |
CH01391/13A CH706859B1 (en) | 2012-08-15 | 2013-08-13 | Gas turbine with arrangement for influencing a temperature gradient and method for operating such a gas turbine. |
CN201320497844.2U CN203669941U (en) | 2012-08-15 | 2013-08-15 | Temperature gradient management device for turbine system |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/586,354 US20140050558A1 (en) | 2012-08-15 | 2012-08-15 | Temperature gradient management arrangement for a turbine system and method of managing a temperature gradient of a turbine system |
Publications (1)
Publication Number | Publication Date |
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US20140050558A1 true US20140050558A1 (en) | 2014-02-20 |
Family
ID=50029676
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US13/586,354 Abandoned US20140050558A1 (en) | 2012-08-15 | 2012-08-15 | Temperature gradient management arrangement for a turbine system and method of managing a temperature gradient of a turbine system |
Country Status (5)
Country | Link |
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US (1) | US20140050558A1 (en) |
JP (1) | JP2014037831A (en) |
CN (1) | CN203669941U (en) |
CH (1) | CH706859B1 (en) |
DE (1) | DE102013108741A1 (en) |
Citations (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5685158A (en) * | 1995-03-31 | 1997-11-11 | General Electric Company | Compressor rotor cooling system for a gas turbine |
US7534087B2 (en) * | 2003-06-16 | 2009-05-19 | Siemens Aktiengesellschaft | Turbomachine, in particular a gas turbine |
Family Cites Families (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20120183398A1 (en) * | 2011-01-13 | 2012-07-19 | General Electric Company | System and method for controlling flow through a rotor |
-
2012
- 2012-08-15 US US13/586,354 patent/US20140050558A1/en not_active Abandoned
-
2013
- 2013-08-09 JP JP2013165704A patent/JP2014037831A/en active Pending
- 2013-08-12 DE DE102013108741.8A patent/DE102013108741A1/en not_active Withdrawn
- 2013-08-13 CH CH01391/13A patent/CH706859B1/en not_active IP Right Cessation
- 2013-08-15 CN CN201320497844.2U patent/CN203669941U/en not_active Expired - Fee Related
Patent Citations (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5685158A (en) * | 1995-03-31 | 1997-11-11 | General Electric Company | Compressor rotor cooling system for a gas turbine |
US7534087B2 (en) * | 2003-06-16 | 2009-05-19 | Siemens Aktiengesellschaft | Turbomachine, in particular a gas turbine |
Also Published As
Publication number | Publication date |
---|---|
CN203669941U (en) | 2014-06-25 |
DE102013108741A1 (en) | 2014-02-20 |
JP2014037831A (en) | 2014-02-27 |
CH706859B1 (en) | 2017-08-15 |
CH706859A2 (en) | 2014-02-28 |
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